Interior cooling circuits in turbine blades

ABSTRACT

A turbine blade that includes an airfoil and a radially extending chamber therein for a coolant. The turbine blade may also include a rib configuration that partitions the chamber into radially extending flow passages. The rib configuration may include: a camber line rib that defines a near-wall flow chamber; and traverse ribs extending between the camber line rib and one of the outer walls so to divide the near-wall flow chamber into successively stacked flow passages that each include a segment of the camber line rib. The segment of the camber line rib of one of the flow passages may have a narrowing profile that includes a profile that narrows from opposing ends toward a neck disposed therebetween.

BACKGROUND OF THE INVENTION

This invention relates to turbine airfoils, and more particularly tohollow turbine airfoils, such as rotor or stator blades, having internalchannels for passing fluids such as air to cool the airfoils.

Combustion or gas turbine engines (hereinafter “gas turbines”) include acompressor, a combustor, and a turbine. As is well known in the art, aircompressed in the compressor is mixed with fuel and ignited in thecombustor and then expanded through the turbine to produce power. Thecomponents within the turbine, particularly the circumferentiallyarrayed rotor and stator blades, are subjected to a hostile environmentcharacterized by the extremely high temperatures and pressures of thecombustion products that are expended therethrough. In order towithstand the repetitive thermal cycling as well as the extremetemperatures and mechanical stresses of this environment, the airfoilsmust have a robust structure and be actively cooled.

As will be appreciated, turbine rotor and stator blades often containinternal passageways or circuits that form a cooling system throughwhich a coolant, typically air bled from the compressor, is circulated.Such cooling circuits are typically formed by internal ribs that providethe required structural support for the airfoil, and include multipleflow paths designed to maintain the airfoil within an acceptabletemperature profile. The air passing through these cooling circuitsoften is vented through film cooling apertures formed on the leadingedge, trailing edge, suction side, and pressure side of the airfoil.

It will be appreciated that the efficiency of gas turbines increases asfiring temperatures rise. Because of this, there is a constant demandfor technological advances that enable turbine blades to withstand everhigher temperatures. These advances sometimes include new materials thatare capable of withstanding the higher temperatures, but just as oftenthey involve improving the internal configuration of the airfoil so toenhance the blades structure and cooling capabilities. However, becausethe use of coolant decreases the efficiency of the engine, newarrangements that rely too heavily on increased levels of coolant usagemerely trade one inefficiency for another. As a result, there continuesto be demand for new airfoil designs that offer internal airfoilconfigurations and coolant circulation that improves coolant efficiency.

A consideration that further complicates design of internally cooledairfoils is the temperature differential that develops during operationbetween the airfoils internal and external structure. That is, becausethey are exposed to the hot gas path, the external walls of the airfoiltypically reside at much higher temperatures during operation than manyof the internal ribs, which, for example, may have coolant flowingthrough passageways defined to each side of them. In fact, a commonairfoil configuration includes a “four-wall” arrangement in whichlengthy inner ribs run parallel to the pressure and suction side outerwalls. It is known that high cooling efficiency can be achieved by thenear-wall flow passages that are formed in the four-wall arrangement,however, the outer walls experience a significantly greater level ofthermal expansion than the inner walls. This imbalanced growth causesstress to develop at the points at which the inner ribs and outer wallsconnect, which may cause low cyclic fatigue that can shorten the life ofthe blade. As such, the development of airfoil structures that usecoolant more efficiently while also reducing stress caused by imbalancedthermal expansion between internal and external regions remains asignificant technological industry objection.

BRIEF DESCRIPTION OF THE INVENTION

The present application thus describes a turbine blade comprising anairfoil defined by outer walls in which a concave shaped pressure sideouter wall and a convex shaped suction side outer wall connect alongleading and trailing edges and, therebetween, form a radially extendingchamber for receiving the flow of a coolant. The turbine blade furtherincludes a rib configuration that partitions the chamber into radiallyextending flow passages. The rib configuration may include: a camberline rib extending alongside one of the outer walls so to define anear-wall flow chamber therebetween; and traverse ribs extending betweenthe camber line rib and the one of the outer walls so to divide thenear-wall flow chamber into successively stacked flow passages, each ofthe flow passages being defined inwardly by a segment of the camber linerib. The segment of the camber line rib of one of the flow passages mayhave a narrowing profile that includes a profile that narrows fromopposing ends toward a neck disposed therebetween.

These and other features of the present application will become apparentupon review of the following detailed description of the preferredembodiments when taken in conjunction with the drawings and the appendedclaims.

BRIEF DESCRIPTION OF THE DRAWINGS

These and other features of this invention will be more completelyunderstood and appreciated by careful study of the following moredetailed description of exemplary embodiments of the invention taken inconjunction with the accompanying drawings, in which:

FIG. 1 is a schematic representation of an exemplary turbine engine inwhich certain embodiments of the present application may be used;

FIG. 2 is a sectional view of the compressor section of the combustionturbine engine of FIG. 1;

FIG. 3 is a sectional view of the turbine section of the combustionturbine engine of FIG. 1;

FIG. 4 is a perspective view of a turbine rotor blade of the type inwhich embodiments of the present invention may be employed;

FIG. 5 is a cross-sectional view of a turbine rotor blade having aninner wall or rib configuration according to conventional design;

FIG. 6 is a cross-sectional view of a turbine rotor blade having aninner wall configuration according to an embodiment of the presentinvention;

FIG. 7 is a cross-sectional view of a turbine rotor blade having aninner wall or rib configuration according to an alternative embodimentof the present invention; and

FIG. 8 is a cross-sectional view of a turbine rotor blade having aninner wall or rib configuration according to an alternative embodimentof the present invention.

DETAILED DESCRIPTION OF THE INVENTION

As an initial matter, in order to clearly describe the current inventionit will become necessary to select certain terminology when referring toand describing relevant machine components within a gas turbine. Whendoing this, if possible, common industry terminology will be used andemployed in a manner consistent with its accepted meaning. Unlessotherwise stated, such terminology should be given a broadinterpretation consistent with the context of the present applicationand the scope of the appended claims. Those of ordinary skill in the artwill appreciate that often a particular component may be referred tousing several different or overlapping terms. What may be describedherein as a being single part may include and be referenced in anothercontext as consisting of multiple components. Alternatively, what may bedescribed herein as including multiple components may be referred toelsewhere as a single part. Accordingly, in understanding the scope ofthe present invention, attention should not only be paid to theterminology and description provided herein, but also to the structure,configuration, function, and/or usage of the component.

In addition, several descriptive terms may be used regularly herein, andit should prove helpful to define these terms at the onset of thissection. These terms and their definitions, unless stated otherwise, areas follows. As used herein, “downstream” and “upstream” are terms thatindicate a direction relative to the flow of a fluid, such as theworking fluid through the turbine engine or, for example, the flow ofair through the combustor or coolant through one of the turbine'scomponent systems. The term “downstream” corresponds to the direction offlow of the fluid, and the term “upstream” refers to the directionopposite to the flow. The terms “forward” and “aft”, without any furtherspecificity, refer to directions, with “forward” referring to the frontor compressor end of the engine, and “aft” referring to the rearward orturbine end of the engine. It is often required to describe parts thatare at differing radial positions with regard to a center axis. The term“radial” refers to movement or position perpendicular to an axis. Incases such as this, if a first component resides closer to the axis thana second component, it will be stated herein that the first component is“radially inward” or “inboard” of the second component. If, on the otherhand, the first component resides further from the axis than the secondcomponent, it may be stated herein that the first component is “radiallyoutward” or “outboard” of the second component. The term “axial” refersto movement or position parallel to an axis. Finally, the term“circumferential” refers to movement or position around an axis. It willbe appreciated that such terms may be applied in relation to the centeraxis of the turbine.

By way of background, referring now to the figures, FIGS. 1 through 4illustrate an exemplary combustion turbine engine in which embodimentsof the present application may be used. It will be understood by thoseskilled in the art that the present invention is not limited to thisparticular type of usage. The present invention may be used incombustion turbine engines, such as those used in power generation,airplanes, as well as other engine types. The examples provided are notmeant to be limiting unless otherwise stated.

FIG. 1 is a schematic representation of a combustion turbine engine 10.In general, combustion turbine engines operate by extracting energy froma pressurized flow of hot gas produced by the combustion of a fuel in astream of compressed air. As illustrated in FIG. 1, combustion turbineengine 10 may be configured with an axial compressor 11 that ismechanically coupled by a common shaft or rotor to a downstream turbinesection or turbine 13, and a combustor 12 positioned between thecompressor 11 and the turbine 13.

FIG. 2 illustrates a view of an exemplary multi-staged axial compressor11 that may be used in the combustion turbine engine of FIG. 1. Asshown, the compressor 11 may include a plurality of stages. Each stagemay include a row of compressor rotor blades 14 followed by a row ofcompressor stator blades 15. Thus, a first stage may include a row ofcompressor rotor blades 14, which rotate about a central shaft, followedby a row of compressor stator blades 15, which remain stationary duringoperation.

FIG. 3 illustrates a partial view of an exemplary turbine section orturbine 13 that may be used in the combustion turbine engine of FIG. 1.The turbine 13 may include a plurality of stages. Three exemplary stagesare illustrated, but more or less stages may be present in the turbine13. A first stage includes a plurality of turbine buckets or turbinerotor blades 16, which rotate about the shaft during operation, and aplurality of nozzles or turbine stator blades 17, which remainstationary during operation. The turbine stator blades 17 generally arecircumferentially spaced one from the other and fixed about the axis ofrotation. The turbine rotor blades 16 may be mounted on a turbine wheel(not shown) for rotation about the shaft (not shown). A second stage ofthe turbine 13 also is illustrated. The second stage similarly includesa plurality of circumferentially spaced turbine stator blades 17followed by a plurality of circumferentially spaced turbine rotor blades16, which are also mounted on a turbine wheel for rotation. A thirdstage also is illustrated, and similarly includes a plurality of turbinestator blades 17 and rotor blades 16. It will be appreciated that theturbine stator blades 17 and turbine rotor blades 16 lie in the hot gaspath of the turbine 13. The direction of flow of the hot gases throughthe hot gas path is indicated by the arrow. As one of ordinary skill inthe art will appreciate, the turbine 13 may have more, or in some casesless, stages than those that are illustrated in FIG. 3. Each additionalstage may include a row of turbine stator blades 17 followed by a row ofturbine rotor blades 16.

In one example of operation, the rotation of compressor rotor blades 14within the axial compressor 11 may compress a flow of air. In thecombustor 12, energy may be released when the compressed air is mixedwith a fuel and ignited. The resulting flow of hot gases from thecombustor 12, which may be referred to as the working fluid, is thendirected over the turbine rotor blades 16, the flow of working fluidinducing the rotation of the turbine rotor blades 16 about the shaft.Thereby, the energy of the flow of working fluid is transformed into themechanical energy of the rotating blades and, because of the connectionbetween the rotor blades and the shaft, the rotating shaft. Themechanical energy of the shaft may then be used to drive the rotation ofthe compressor rotor blades 14, such that the necessary supply ofcompressed air is produced, and also, for example, a generator toproduce electricity.

FIG. 4 is a perspective view of a turbine rotor blade 16 of the type inwhich embodiments of the present invention may be employed. The turbinerotor blade 16 includes a root 21 by which the rotor blade 16 attachesto a rotor disc. The root 21 may include a dovetail configured formounting in a corresponding dovetail slot in the perimeter of the rotordisc. The root 21 may further include a shank that extends between thedovetail and a platform 24, which is disposed at the junction of theairfoil 25 and the root 21 and defines a portion of the inboard boundaryof the flow path through the turbine 13. It will be appreciated that theairfoil 25 is the active component of the rotor blade 16 that interceptsthe flow of working fluid and induces the rotor disc to rotate. Whilethe blade of this example is a turbine rotor blade 16, it will beappreciated that the present invention also may be applied to othertypes of blades within the turbine engine 10, including turbine statorblades 17. It will be seen that the airfoil 25 of the rotor blade 16includes a concave pressure side outer wall 26 and a circumferentiallyor laterally opposite convex suction side outer wall 27 extendingaxially between opposite leading and trailing edges 28, 29 respectively.The sidewalls 26 and 27 also extend in the radial direction from theplatform 24 to an outboard tip 31. (It will be appreciated that theapplication of the present invention may not be limited to turbine rotorblades, but may also be applicable to stator blades. The usage of rotorblades in the several embodiments described herein is exemplary unlessotherwise stated.)

FIG. 5 shows an internal wall construction as may be found in a rotorblade airfoil 25 having a conventional design. As indicated, the outersurface of the airfoil 25 may be defined by a relatively thin pressureside outer wall 26 and suction side outer wall 27, which may beconnected via a plurality of radially extending and intersecting ribs60. The ribs 60 are configured to provide structural support to theairfoil 25, while also defining a plurality of radially extending andsubstantially separated flow passages 40. Typically the ribs 60 extendradially so to partition the flow passages 40 over much of the radialheight of the airfoil 25, but, as discussed more below, the flow passagemay be connected along the periphery of the airfoil so to define acooling circuit. That is, the flow passages 40 may fluidly communicateat the outboard or inboard edges of the airfoil 25, as well as via anumber of smaller crossover passages or impingement apertures (notshown) that may be positioned therebetween. In this manner certain ofthe flow passages 40 together may form a winding or serpentine coolingcircuit. Additionally, film cooling ports (not shown) may be includedthat provide outlets through which coolant is released from the flowpassages 40 onto the outer surface of the airfoil 25.

The ribs 60 may include two different types, which then, as providedherein, may be subdivided further. A first type, a camber line rib 62,is typically a lengthy rib that extends in parallel or approximatelyparallel to the camber line of the airfoil, which is a reference linestretching from the leading edge 28 to the trailing edge 29 thatconnects the midpoints between the pressure side outer wall 26 and thesuction side outer wall 27. As is often the case, the conventionalconfiguration of FIG. 5 includes two camber line ribs 62, a pressureside camber line rib 63, which also may be referred to as the pressureside inner wall given the manner in which it is offset from and close tothe pressure side outer wall 26, and a suction side camber line rib 64,which also may be referred to as the suction side inner wall given themanner in which it is offset from and close to the suction side outerwall 27. As mentioned, this type of design is often referred to ashaving a “four-wall” configuration due to the prevalent four main wallsthat include the two sidewalls 26, 27 and the two camber line ribs 63,64.

The second type of rib is referred to herein as a traverse rib 66.Traverse ribs 66 are the shorter ribs that are shown connecting thewalls and inner ribs of the four-wall configuration. As indicated, thefour walls may be connected by a number of the traverse ribs 66, whichmay be further classified according to which of the walls each connects.As used herein, the traverse ribs 66 that connect the pressure sideouter wall 26 to the pressure side camber line rib 63 are referred to aspressure side traverse ribs 67. The traverse ribs 66 that connect thesuction side outer wall 27 to the suction side camber line rib 64 arereferred to as suction side traverse ribs 68. Finally, the traverse ribs66 that connect the pressure side camber line rib 63 to the suction sidecamber line rib 64 are referred to as center traverse ribs 69.

In general, the purpose of four-wall internal configuration in anairfoil 25 is to provide efficient near-wall cooling, in which thecooling air flows in channels adjacent to the outer walls 26, 27 of theairfoil 25. It will be appreciated that near-wall cooling isadvantageous because the cooling air is in close proximity of the hotouter surfaces of the airfoil, and the resulting heat transfercoefficients are high due to the high flow velocity achieved byrestricting the flow through narrow channels. However, such designs areprone to experiencing low cycle fatigue due to differing levels ofthermal expansion experienced within the airfoil 25, which, ultimately,may shorten the life of the rotor blade. For example, in operation, thesuction side outer walls 27 thermally expands more than the suction sidecamber line rib 64. This differential expansion tends to increase thelength of the camber line of the airfoil 25, and, thereby, causes stressbetween each of these structures as well as those structures thatconnect them. In addition, the pressure side outer wall 26 alsothermally expands more than the cooler pressure side camber line rib 63.In this case, the differential tends to decrease the length of thecamber line of the airfoil 25, and, thereby, cause stress between eachof these structures as well as those structures that connect them. Theoppositional forces within the airfoil that, in the one case, tends todecrease the airfoil camber line and, in the other, increase it, canlead to further stress concentrations. The various ways in which theseforces manifest themselves given an airfoil's particular structuralconfiguration and the manner in which the forces are then balanced andcompensated for becomes a significant determiner of the part life of therotor blade 16.

More specifically, in a common scenario, the suction side outer wall 27tends to bow outward at the apex of its curvature as exposure to thehigh temperatures of the hot gas path cause it to thermally expand. Itwill be appreciated that the suction side camber line rib 64, being aninternal wall, does not experience the same level of thermal expansionand, therefore, does not have the same tendency to bow outward. Thecamber line rib 64 then resists the thermal growth of the outer wall 27.Because conventional designs have camber line ribs 62 formed with stiffgeometries that provide little or no compliance, this resistance and thestress concentrations that result from it can be substantial.Exacerbating the problem, the traverse ribs 66 used to connect thecamber line rib 62 to the outer wall 27 are formed with linear profilesand generally oriented at right angles in relation to the walls thatthey connect. This being the case, the traverse ribs 66 operate tobasically hold fast the “cold” spatial relationship between the outerwall 27 and the camber line rib 64 as the heated structures expand atsignificantly different rates. Accordingly, with little or no “give”built into the structure, conventional arrangements are ill-suited atdefusing the stress that concentrates in certain regions of thestructure. The differential thermal expansion bus results in low cyclefatigue issues that shorten component life.

Many different internal airfoil cooling systems and structuralconfigurations have been evaluated in the past, and attempts have beenmade to rectify this issue. One such approach proposes overcooling theouter walls 26, 27 so that the temperature differential and, thereby,the thermal growth differential are reduced. It will be appreciated,though, that the way in which this is typically accomplished is toincrease the amount of coolant circulated through the airfoil. Becausecoolant is typically air bled from the compressor, its increased usagehas a negative impact on the efficiency of the engine and, thus, is asolution that is preferably avoided. Other solutions have proposed theuse of improved fabrication methods and/or more intricate internalcooling configurations that use the same amount of coolant, but use itmore efficiently. While these solutions have proven somewhat effective,each brings additional cost to either the operation of the engine or themanufacture of the part, and does nothing to directly address the rootproblem, which is the geometrical deficiencies of conventional design inlight of how airfoils grow thermally during operation.

The present invention generally teaches certain curving or bubbled orsinusoidal or wavy internal ribs (hereinafter “wavy ribs”) thatalleviate imbalanced thermal stresses that often occur in the airfoil ofturbine blades. Within this general idea, the present applicationdescribes several ways in which this may be accomplished, which includewavy camber line ribs 62 and/or traverse ribs 66, as well as certaintypes of angled connections therebetween, and camber line ribs 62 thathave a narrowing profile between connecting traverse ribs 66. It will beappreciated that these novel configurations—which, as delineated in theappended claims, may be employed separately or in combination—reduce thestiffness of the internal structure of the airfoil 25 so to providetargeted flexibility by which stress concentrations are dispersed andstrain off-loaded to other structural regions that are better able towithstand it. This may include, for example, off-loading to a regionthat spreads the strain over a larger area, or, perhaps, structure thatoffloads tensile stress for a compressive load, which is typically morepreferable. In this manner, life-shortening stress concentrations andstrain may be avoided.

FIGS. 6 through 8 provide cross-sectional views of a turbine rotor blade16 having an inner wall configuration according to certain aspects ofthe present invention. The present invention involves the configurationof ribs 60 that are used as both structural support as well aspartitions that divide hollow airfoils 25 into interconnected radiallyextending flow passages 40. These flow passages 40 direct a flow ofcoolant through the airfoil 25 in a particular manner so that its usageis targeted and more efficient. Though the examples provided herein areshown as they might be used in a turbine rotor blades 16, it will beappreciated that the same concepts also may be employed in turbinestator blades 17. FIG. 6 illustrates a rib configuration of the presentinvention that has a camber line rib 62 having a wavy profile. (As usedherein, the term “profile” is intended to refer to the shape the ribshave in the cross-sectional views of FIGS. 6 through 8.) A camber linerib 62, as described above, is one of the longer ribs that extends froma position near the leading edge 28 of the airfoil 25 toward thetrailing edge 29. These ribs are referred to as “camber line ribs”because the path they trace is approximately parallel to the camber lineof the airfoil 25, which is a reference line extending between theleading edge 28 and the trailing edge 29 of the airfoil 25 through acollection of points that are equidistant between the concave pressureside outer wall 26 and the convex suction side outer wall 27. Accordingto the present application, a “wavy profile” includes one that isnoticeably curved and sinusoidal in shape, as indicated. In other words,the “wavy profile” is one that presents a back-and-forth “S” profile.

The segment or length of the camber line rib 62 that is configured withthe wavy profile may vary depending on design criteria. In the providedexamples the wavy camber line rib 62 typically stretches from a positionnear the leading edge 28 of the airfoil 25 to a position that is beyondthe midpoint of the camber line of the airfoil 25. It will beappreciated that the wavy portion of the camber line rib 62 may beshorter in length while still providing the same types of performanceadvantages discussed herein. The number of curves as well as the lengthof the wavy segment of the camber line rib 62 may be varied to achievethe best results. In certain embodiments, the wavy camber line rib 62 ofthe present invention is defined by the number of completeback-and-forth “S” shapes it contains. In a preferred embodiment of thistype, the wavy camber line rib 62 includes at least one continuousback-and-forth “S” shape. In another embodiment, the wavy camber linerib 62 includes at least two consecutive and continuous back-and-forth“S” shapes. It will be appreciated that the examples provided in FIGS. 6and 7 each trace paths having more than two full “S” shapes. In regardto overall length, the wavy segment of the camber line rib 62 may extendfor a substantial portion of the length of the camber line of theairfoil 25. For example, as shown in FIGS. 6 and 7, in a preferredembodiment, the wavy portion of the camber line rib 62 is at least 69%of the length of the camber line of the airfoil 25. In other words, thewavy portion of the camber line rib 62 originates near the leading edge28 of the airfoil 25 and extend rearward and well beyond the apex of thecurvature of the airfoil 25.

It will be appreciated that, given its winding profile, a wavy camberline rib 62 traces a path that varies in its directional heading. Thewavy camber line rib 62 of the present invention may still be describedas having a general arcing path across which it winds, and that thispath typically extends from an origination point near the leading edge28 and a trailing point near the trailing edge 29 of the airfoil 25. Itwill be appreciated that, in the case of a wavy camber line rib 62, itis this general arcing path that is roughly parallel to the camber lineof the airfoil 25.

Many known airfoil 25 configurations, such as the four-wall example ofFIG. 5 discussed above, include two camber line ribs 62. This type ofconfiguration may be described as having a pressure side camber line rib63 that resides nearer the pressure side outer wall 26, and a suctionside camber line rib 64 that resides nearer the suction side outer wall27. The present invention, as shown in FIGS. 6 and 7, may includeconfigurations in which both the suction side camber line rib 64 and thepressure side camber line rib 63 are formed as wavy ribs. In alternativeembodiments, only one of these camber line ribs 62 may have a wavyprofile. It will be appreciated that the present invention may also beemployed in configurations having only a single camber line rib 62.

In airfoils 25 that include two camber line ribs 62, it will beappreciated that the pressure side camber line rib 63 and the suctionside camber line rib 64 define a center flow passage 40. The wavyprofile for each of the pressure side camber line rib 63 and the suctionside camber line rib 64 may be defined relative to the shape taken bysuccessive segments of the camber line rib 62 facing center flow passage40. That is, for example, relative to the central flow passage 40, thewavy profile of the camber line rib 62 may be described as including twosuccessive segments in which a first concave segment transitions to asecond convex segment. In an alternative embodiment, the wavy profilemay include four successive segments in which: a first concave segmenttransitions to a second convex segment; the second convex segmenttransitions to a third concave segment; and the third concave segmenttransitions to a fourth convex segment.

In certain embodiments, the present invention teaches certainconfigurations of traverse ribs 66 that may be employed to tune thecompliancy of the airfoil 25. As used herein, traverse ribs 66 are theshorter ribs that extend across the airfoil 25. Traverse ribs 66 areused to connect camber line ribs 62 to either other camber line ribs orone of the outer walls 26, 27 of the airfoil 25. It will be appreciatedthat, configured in this way, traverse ribs 66 also serve as partitionsto the flow passages 40 formed between the outer walls 26, 27 and thecamber line ribs 62. As illustrated, the pressure side outer wall 26 andthe pressure side camber line rib 63 are configured to define a pressureside flow passage 40 between them. Similarly, the suction side outerwall 27 and the suction side camber line rib 64 are configured to definea suction side flow passage 40 between them. Between the suction sidecamber line rib 64 and the pressure side camber line rib 63, the centerflow passage 40 is defined. As indicated, these flow passages 40 thenmay be subdivided by the traverse ribs 66. In certain embodiments of thepresent invention, several pressure side traverse ribs 67 connect thepressure side outer wall 26 to the pressure side camber line rib 63.Thusly, the pressure side traverse ribs 67 divide the pressure side flowpassage 40 into a number of separate, axially-stacked flow passages 40.Likewise, several suction side traverse ribs 68 connect the suction sideouter wall 27 to the suction side camber line rib 64 and divide thesuction side flow passage 40 into a number of separate, axially-stackedflow passages 40. Center traverse ribs 69 connect the pressure sidecamber line rib 63 to the suction side camber line rib 64, and similarlypartition the center flow passage.

The camber line 62 ribs and traverse ribs 66 may be configured asradially extending walls. That is, these ribs may form the profilesshown in the cross-sectional views of FIGS. 6 through 8 while extendingradially between the two ends of the airfoil 25. In this manner, thepressure side flow passages, suction side flow passages, and center flowpassages 40 may extend radially between an inboard end that is near theinterface between the airfoil 25 and the blade root 21 and an outboardend that is near the outboard tip 31 of the airfoil 25. In usage, asupply of coolant may be delivered to one or more of the inboard ends ofthe flow passages 40 via a supply passage that extends through a bladeroot 21. It will be appreciated that the flow passages 40 may beselectively connected at their inboard or outboard ends so to create aserpentine coolant path through the airfoil 25.

As illustrated in FIG. 6, rib configurations of the present inventionmay include several traverse ribs 66 on each of the pressure and suctionsides of the airfoil 25. In a preferred embodiment, at least fivepressure side traverse ribs 67 and five suction side traverse ribs 68may be included. Multiple center traverse ribs 69 also may be provided,though a single traverse rib 69 also may be used in other embodiments.As shown, in preferred embodiments, the present invention may include atleast two center traverse ribs 69. The present invention furtherdescribes a connection assembly by which the traverse ribs 66 connect tothe outer walls 26, 27 and/or the camber line ribs 62. It will beappreciated that the angle in which traverse ribs 66 intersect suchwalls 26, 27, 62 may be described by an “angle of connection”. (It willbe appreciated that the “angle of connection” referred to is the smallerof the two angles formed on both sides of each end of a traverse ribbetween the traverse rib and the wall it intersects.) In conventionalairfoil configurations, as mentioned above, the angle of connection is asteep one, generally being close to 90°. It will be appreciated thatsteep angles like this make for a stiff structure. The present inventionteaches angles that are significantly less than 90° as a way in whichthe airfoil 25 structure, or targeted areas of the structure, may bemade more compliant. According to one embodiment, as shown in FIGS. 6and 7, at least two of the pressure side traverse ribs 67 may beconfigured so to have an angle of connection with the pressure sideouter wall 26 of less than 60 degrees. According to another embodiments,as indicated, at least two of the suction side traverse ribs 68 may beconfigured so to form an angle of connection with the suction side outerwall 27 of less than 60 degrees. The center traverse ribs 69 may besimilarly formed, and configurations of the present invention includehaving at least one angle of connection of less than 60 degrees at eachof the suction side camber line rib 64 and the pressure side camber linerib 63. Where greater compliance is required, embodiments may includehaving three of the pressure side traverse ribs 67 and three of thesuction side traverse ribs 68 configured so to have an angle ofconnection with the outer walls 26, 27 of less than 60 degrees, and atleast two of the center traverse ribs 69 may be configured so to form anangle of connection of less than 60 degrees at each of the suction sidecamber line rib 64 and the pressure side camber line rib 63.

The present invention further describes another manner in which traverseribs 66 may enhance structural compliancy. Traverse ribs 66 typicallyare formed having a linear profile, which, as will be appreciated,results in a stiff and unyielding configuration. Pursuant to certainembodiments of the present invention, traverse ribs 66 are configuredhaving a curved profile. Specifically, as shown in each of the examplesin FIGS. 6 through 8, the center traverse ribs 69 may include an curved,arcuate or arcing profile. With this profile, the traverse ribs 66become much more compliant and able to accommodate relative movementbetween the structural walls that they connect. The direction in whichthe curved arcing profile of the traverse rib is oriented may bemanipulated so to accommodate the types of expected stresses. Accordingto one preferred embodiment, as illustrated in FIG. 6, the arc of thecenter traverse rib 69 may be directed such that the concave face of thecenter traverse ribs 69 is directed toward the leading edge 28 of theairfoil 25. This orientation may be done to all of the center traverseribs 69 that are included in a particular structure or a fractionthereof. In an alternative embodiment, as illustrated in FIG. 7, the arcof the center traverse ribs 69 may be directed such that the convex faceof the traverse rib is directed away from the leading edge 28 of theairfoil 25. This type of profile may be used on all of center traverseribs 69 or only a fraction of them.

According to the present invention, the internal structure of an airfoilmay include wavy ribs along the camber line direction of the airfoil. Bymaking the camber line rib 62 into a spring in this way, the internalbackbone of the airfoil may be made more compliant so that performanceadvantages may be achieved. In addition, the traverse ribs of theairfoil structure may be curved so to further soften the load path, aswell as making more compliant connections with the ribs 62 and outerwalls 26, 27 that they connect. Whereas standard linear rib designsexperience high stress and low cyclic life due to the thermal fightbetween the internal cooling cavity walls and the much hotter outerwalls, the present invention provides a spring-like construction that isbetter able to disburse stress concentrations, which, as providedherein, may be used to improve the life of the component.

FIGS. 7 and 8 illustrate another aspect of the present invention inwhich one or more segments 73 of a camber line rib 62 or formed having anarrowing profile 72. According to the present application and as usedherein, a “narrowing profile” is one in which rib thickness narrows fromopposing ends 75 of the segment 73 so to form a neck or narrow regionbetween them. Describe another way, a narrowing profile 72 is one inwhich the camber line rib segment 73 has a thickness that graduallyflares or thickens in both directions as it extends 75 outward from theneck 74. It will be appreciated that, given this configuration, thenarrowing profile 72 is similar to an hourglass shape.

The narrowing profile 72 of the present invention may be included inflow passages 40 that are inwardly defined by a camber line rib 62. Inpreferred embodiments, as indicated, the narrowing profile 72 is appliedto a camber line rib segment 73 that is defined in relation to theconfiguration of near wall flow passages 40 formed between the camberline rib 62 and one of the outer walls 26, 27 of the airfoil 25. Inpreferred embodiments, the segment 73 having the narrowing profile 72 isdefined between successive traverse ribs 66. Specifically, it will beappreciated that the near wall flow passages are defined between twosuccessive traverse ribs 66 that connect the outer wall 26, 27 of theairfoil with five to a camber line rib 62. In preferred embodiments, thecamber line rib segment 73 to which the narrowing profile 72 is appliedis the length of camber line rib 62 defined between the locations atwhich each of those successive traverse ribs 66 intersect the camberline rib 62. Defined in this manner, the camber line rib segment 73 maybe referred to as a traverse rib-to-traverse rib segment.

As indicated in FIG. 8, the narrowing profile 72 may be used with camberline ribs 62 that are wavy or sinusoidal in form, in which case, it maybe used with any of the configurations that are discussed above inregard to the other aspects of the present invention. The narrowingprofile 72, however, is not limited to this type of usage. As shown inFIG. 7, the narrowing profile 72 may be used in conjunction withtraditionally formed camber line ribs 62, i.e., those having an axisthat is substantially linear. In preferred embodiments, the narrowingprofile 72 may be used on the camber line ribs 62 of two or moreconsecutively formed flow passages 40. It will be appreciated that,though, the narrowing profile 72 may still offer certain performanceadvantages if used on the camber line segment 73 of a single flowpassage 40. It will be appreciated also that the narrowing profile 72may be used on camber line segments 73 of pressure side camber line ribs62, suction side camber line ribs 62, or both.

As provided in FIGS. 7 and 8, the narrowing profile 72 may include acurved or linear shape. In a preferred embodiment, the narrowing profile72 includes a contoured shape that gradually widens from the neck 74. Asindicated, in a preferred embodiment, the neck 74 may be positioned ator near the midpoint of the segment 73 of the camber line rib 62. Itwill be understood that the neck 74 may be described as having a “neckthickness”, while each of the opposing ends 75 of the segment 73 may bedescribed as having an “end thickness”, which, respectively, arerepresented in FIGS. 7 and 8 as distance 77 and distance 78. Inexemplary embodiments, the end thickness 78 of each of the ends 75 is atleast 1.5 times greater than the neck thickness 77. In an alternativeembodiment, the end thickness 78 of each of the ends 75 is at least 2times greater than the neck thickness 77. The narrowing profile betweeneach of the opposing ends 75 and the neck 74 may narrow in a smoothand/or constant manner, which will minimize stress concentrations.

It will be appreciated that the narrowing profile 72 of the presentinvention provides another manner in which the compliancy of the airfoilstructure may be enhanced or further tuned so to minimize or spreadstresses resulting from imbalanced thermal growth. In a preferredembodiment, the narrowing profile 72 may be used in conjunction withaspects of the wavy camber line ribs 62 to enhance the compliancy of the“spring” these ribs generally form. In this case, it will be understoodthat the narrowing profile 72 enhances the spring effect of the wavywalls and, thereby, reduces the stress experienced by the internalstructure of the airfoil 25. Along these lines, with further referenceto FIG. 8, it will be appreciated the narrowing profile 72 provides thecurved segments 73 of the wavy ribs the compliancy to straighten inresponse to a tensile load, and further bend in response to acompressive one. As stated, the narrowing profile 72 may also be usedwith more conventional linear type camber line ribs 62 as a means ofproviding at least some level of enhanced compliancy to the airfoilstructure. The resulting reduced stress levels that these type of ribconfigurations provide may be used to extend component cyclic life.

As one of ordinary skill in the art will appreciate, the many varyingfeatures and configurations described above in relation to the severalexemplary embodiments may be further selectively applied to form theother possible embodiments of the present invention. For the sake ofbrevity and taking into account the abilities of one of ordinary skillin the art, all of the possible iterations is not provided or discussedin detail, though all combinations and possible embodiments embraced bythe several claims below or otherwise are intended to be part of theinstant application. In addition, from the above description of severalexemplary embodiments of the invention, those skilled in the art willperceive improvements, changes and modifications. Such improvements,changes and modifications within the skill of the art are also intendedto be covered by the appended claims. Further, it should be apparentthat the foregoing relates only to the described embodiments of thepresent application and that numerous changes and modifications may bemade herein without departing from the spirit and scope of theapplication as defined by the following claims and the equivalentsthereof.

We claim:
 1. A turbine blade comprising an airfoil defined by outerwalls in which a concave shaped pressure side outer wall and a convexshaped suction side outer wall connect along leading and trailing edgesand, therebetween, form a radially extending chamber for receiving aflow of a coolant, the turbine blade further comprising: a ribconfiguration that partitions the chamber into radially extending flowpassages, wherein the rib configuration includes: a camber line ribextending alongside one of the outer walls so to define a near-wall flowchamber therebetween; and traverse ribs extending between the camberline rib and the one of the outer walls so to divide the near-wall flowchamber into successively stacked flow passages, each of the flowpassages being defined inwardly by a segment of the camber line rib;wherein the segment of the camber line rib of one of the flow passagescomprises a narrowing profile, the narrowing profile comprising aprofile that narrows from opposing ends toward a neck disposedtherebetween; wherein the narrowing profile comprises a cross-sectionalprofile; wherein the segment of the camber line rib comprises a traverserib-to-traverse rib segment in which each of the opposing ends aredefined at an intersection with one of the traverse ribs; wherein thetraverse ribs form at least two axially stacked flow passages, andwherein the segment of the camber line rib for each of the two axiallystacked flow passages each comprises the narrowing profile; and whereineach of the opposing ends comprises an end thickness, and the neckcomprises a neck thickness, and wherein the end thickness is at least1.5 times greater than the neck thickness.
 2. The turbine bladeaccording to claim 1, wherein the traverse ribs form three successivelystacked flow passages; and wherein the segment of the camber line ribfor each of the three successively stacked flow passages each comprisesthe narrowing profile.
 3. The turbine blade according to claim 2,wherein the segment of the camber line rib for each of the threesuccessively stacked flow passages each comprises a linear axis.
 4. Theturbine blade according to claim 2, wherein the segments of the camberline rib for the three successively stacked flow passages comprise awavy axis.
 5. The turbine blade according to claim 4, wherein the wavyaxis comprises at least one back-and-forth “S” shape.
 6. The turbineblade according to claim 2, wherein the camber line rib comprises apressure side camber line rib that extends alongside and near thepressure side outer wall; wherein the traverse ribs extend between thepressure side camber line rib and the pressure side outer wall to formthe three successively stacked flow passages, each of the flow passagesbeing defined by the pressure side outer wall, two opposing traverseribs, and, opposite the pressure side outer wall, the segment of thepressure side camber line rib.
 7. The turbine blade according to claim6, wherein the pressure side camber line rib comprises a wavy profilethat includes at least two consecutive back-and-forth “S” shapes.
 8. Theturbine blade according to claim 2, wherein the camber line ribcomprises a suction side camber line rib that extends alongside and nearthe suction side outer wall; wherein the traverse ribs extend betweenthe suction side camber line rib and the suction side outer wall to formthe three successively stacked flow passages, each of the flow passagesbeing defined by the suction side outer wall, two opposing traverseribs, and, opposite the suction side outer wall, the segment of thesuction side camber line rib.
 9. The turbine blade according to claim 8,wherein the suction side camber line rib comprises a wavy profile thatincludes at least two consecutive back-and-forth “S” shapes.
 10. Theturbine blade according to claim 1, wherein the rib configurationincludes two camber line ribs: a pressure side camber line rib thatextends alongside the pressure side outer wall and forms a near-wallflow chamber therebetween, and a suction side camber line rib thatextends alongside the suction side outer wall and forms a near-wall flowchamber therebetween; wherein the traverse ribs extend between thepressure side camber line rib and the pressure side outer wall to format least two successively stacked flow passages, wherein the segment ofthe pressure side camber line rib for each of the two successivelystacked flow passages each comprises the narrowing profile; and whereinthe traverse ribs extend between the suction side camber line rib andthe suction side outer wall to form at least two successively stackedflow passages, wherein the segment of the suction side camber line ribfor each of the two successively stacked flow passages each comprisesthe narrowing profile.
 11. The turbine blade according to claim 10,wherein at least one of the pressure side camber line rib and thesuction side camber line rib comprises a wavy profile that includes atleast two consecutive back-and-forth “S” shapes.
 12. The turbine bladeaccording to claim 1, wherein the neck of the narrowing profile isdisposed near a midpoint between the opposing ends of the segment of thecamber line rib.
 13. The turbine blade according to claim 1, wherein theend thickness is at least 2 times greater than the neck thickness. 14.The turbine blade according to claim 1, wherein the narrowing betweeneach of the opposing ends and the neck is constant.
 15. The turbineblade according to claim 1, wherein the narrowing between each of theopposing ends and the neck is smooth.
 16. The turbine blade according toclaim 1, wherein the turbine blade comprises a turbine rotor blade.